JP4065409B2 - Resolvable thrust vector control wing - Google Patents

Abstract

A dissolvable jet vane (22/30) is a composite structure, having a support frame (38), a plug leading edge (40) connected to the forward edge (42) of the frame (38), and an insulation layer (44) on the side walls of the support frame (38). The dissolvable jet vane materials withstand the pressure and thermal loads associated with missile steering during the first few seconds of rocket boost until the missile attains sufficient speed to use conventional external aerodynamic control fins for steering control. Once control passes to the external fins, the jet vanes rapidly and uniformly dissolve in the exhaust stream. The dissolvable jet vane provides a lightweight, reliable means of removing steering jet vanes from the exhaust stream of a solid rocket motor nozzle.

Description

  The present invention relates to a missile control system, and more particularly to a thrust vector control wing and system attached to the aft of a missile used for maneuvering the missile during launch, and a method of making and using such a control wing.

This invention is made with US Government support under contract number DASG60-98-C-0061 awarded by the US Department of Defense. The US government has certain rights in this invention.
Attack missiles, including aircraft and some cruise missiles, often fly at low altitudes, for example, to avoid detection by enemy radar. In such a situation, the target, such as a ship, has only a few minutes to identify the threat posed by the incoming aircraft or missile and take countermeasures such as shooting a defense missile. Therefore, it is desirable to position defense missiles so that attack missiles become effective very quickly.

  Ground or ship-borne defense missiles are typically launched from a substantially vertically oriented cylinder (canister). Missiles typically maneuver control systems that include external aerodynamic control surfaces for missile guidance. Before the aerodynamic control surface or fins can affect any large mobility, the missile is here aerodynamic control speed and the aerodynamic control speed to allow sufficient air to flow over the aerodynamic control surface. You have to reach a certain minimum speed to call. In a ballistic launch trajectory, the missile reaches thousands of feet before the aerodynamic control surface pitches the missile and begins to search for the incoming missile threat. As a result, the ballistic launch trajectory is not efficient and time consuming, limiting the missile sensor's line-of-sight capability for optimal target detection and tracking.

  A number of systems have been developed to reduce response time after sensing threats that are maneuvering missiles before reaching aerodynamic control speed. However, current state-of-the-art devices are usually acceptable for some uses, but have been found to be inappropriate for many applications. For example, a removable jet tab system formed with an auxiliary propulsion device attached to a missile fin is incompatible with a foldable control surface. Folding control surfaces are necessary for missiles to be installed in launch canisters that usually have strict volume limitations. The removable jet tab system requires an increase in the cross-sectional area of the launch canister due to the additional volume occupied by the jet tab external to the missile fuselage.

  Existing systems can be classified as either non-removable or removable, the latter often having redundant control electronics. Non-removable systems limit range and performance due to rocket thrust degradation through missile trajectories. A removable self-biasing mechanism is heavier and more complex than a non-removable system. Increased complexity reduces reliability, and the addition of weight requires more missile launches and rocket thrust to fly to the target. Actively removable systems typically use fired fired firing mechanisms and releasable power-coupled drives, which incurs weight, complexity, and risk of numerous operational failures. Furthermore, the action of injecting the control system removes the missile from its intended trajectory.

  In order to overcome the shortcomings of conventional systems, systems have been developed that place mechanisms in the rocket engine exhaust plume for control purposes, with immediate control upon launch. Usually, the objective is to pitch over the missile (rotating the missile around an axis that intersects the longitudinal axis and the previous flight direction during launch), and rolling usually interferes with the operation of the missile guidance system and the control surface. This is a problem that is minimized at low speeds by positioning it in the exhaust jet.

  So-called “corrosive” control surfaces have been developed that are placed in the exhaust gas passages of rocket engines and separate after a period of time. However, frequent separation of these larger than normal pieces is unacceptable. Therefore, it is desirable to avoid discharging large pieces of material from missiles during launch and flight.

  The present invention provides a thrust vector controlled jet wing, a jet wing control system, a missile having such a jet wing, and a method of manufacturing such a jet wing. Jet wings form an aerodynamic surface that generates vehicle maneuvering force by deflecting the propulsion jet during missile launch. When the aerodynamic control speed is reached and the aerodynamic control surface external to the vehicle body has command authority, the jet wing breaks down into granular particles in the propulsion jet. Therefore, the propulsion efficiency or special impulse of the rocket motor does not degrade more than needed to perform the launch operation, the response time is reduced, and the missile is in the aerodynamic control phase of the energized flight to the target. Inside, improved athletic performance is shown.

  Decomposable jet control wings offer many advantages in missile design, structure and / or performance. For example, a decomposable control wing eliminates or minimizes many tedious implementations to determine the desired minimum inefficiency for the entire missile system. The control wing breaks down properly on a time basis and provides a releasable control wing for manipulating the missile to control pitch, yaw and rolling immediately upon launch. The rocket motor propulsion efficiency or special impulse is not degraded more than necessary to perform the launch operation, and thus the missile exhibits improved kinematic performance during the energized flight phase to the target. The interception of highly mobile targets at extended distances is further enhanced by the releasable abandonment of the control wing after pitch over. The ease of control of thrust blades without bias by the firework-biased injection mechanism and the greater reliability gained from the simplicity of the system are additional benefits that reduce cost and risk.

  Decomposable control wings are possible through the use of a number of advanced and complex materials designed to perform different individual functions based on time limitations and take advantage of known high temperature performance characteristics and environmental limitations Are integrated or laminated together to achieve a combined pre-programmed structural capability. As a result, the failure of complex control wings manufactured in accordance with the present invention can be accurately controlled in ways that could not be foreseen with conventional “corrosion” material designs. Complex resolvable jet wings also provide an inexpensive and processable thrust vector control method that retrofits high performance missiles for low speed surface launch applications due to thrust vector requirements.

  The decomposable jet wing provides a lightweight and reliable means of removing the steering jet wing from the exhaust stream of a solid rocket motor nozzle. The decomposable jet wing material uses the pressure associated with the missile maneuver for the first few seconds of the rocket boost until it acquires sufficient speed using a normal external aerodynamic control surface for missile maneuver control. And withstand thermal loads. When control is transferred to the external fins, the jet vanes are quickly and uniformly decomposed in the exhaust stream.

  According to one feature of the invention, the decomposable thrust vector control wing includes a frame and a thermal protection layer on at least a portion of the frame. According to an embodiment of the present invention, the dismantleable control wing further includes a corrosion protection material on at least the front edge of the frame. This anticorrosion material forms an insert attached to the front edge of the frame, which insert has a carbon-carbon structure and a surface coating selected from the group comprising ceramic, carbide, and metallic materials on the structure. And the surface coating comprises one or more materials selected from the group comprising hafnium boride ceramic, zirconium boride ceramic, hafnium carbide, tantalum carbide or metal rhenium refractory coating, and the surface coating is typically about 5,000. The insert has a thickness less than one inch (0.127 mm) thickness, and the insert includes one or more materials selected from ceramic and oxide-based oxygen permeable and volatile barriers; The insert usually has a T-shaped cross section with a T tail connected to the frame, and / or the insert is usually a frame Having a U-shaped cross section having an open end attached U-shaped.

  According to one or more embodiments of the present invention, the frame includes an organic resin reinforced with graphite, the organic resin reinforced with graphite is a PMR-15 based polyimide, PT cyanate ester, bismarimide, phthalonitrile, A high temperature organic resin reinforced with carbon containing one or more resins selected from the group containing Avimid N, the thermal protection layer contains an ablation insulator, and the thermal protection layer is cut Including one or more materials selected from the group comprising phenolic resin reinforced with quartz, carbon and silica fibers, the control assembly includes a control shaft connected to the frame, the control shaft being fire resistant Formed from material and / or the frame is substantially planar and the control wings pass through the frame surface. And further comprising an integrated blast shield extending in a plane perpendicular.

In yet another embodiment, the disassembleable control wing includes a control assembly connected to the frame to control the direction of the wing.
In accordance with another aspect of the present invention, a propulsion vector control system for a missile comprising a motor for propelling a missile that generates an exhaust jet includes a resolvable control wing and a control assembly for controlling the direction of the control wing. Including. The system can be attached to a missile so that the control vanes are located in the passage of the exhaust jet.

According to another aspect of the present invention, the missile includes a motor for propelling the missile that generates the exhaust jet and a decomposable control vane mounted in the passage of the exhaust jet.
According to yet another aspect of the present invention, a method for maneuvering a missile during launch includes placing a movable control wing in the missile exhaust jet path, launching the missile, and controlling the wing to maneuver the missile. Includes moving steps as possible. The missile launch step ignites the motor to discharge the exhaust jet and disassembles the wings in the exhaust jet within a predetermined time period.

According to one or more embodiments of the present invention, the step of moving the control wing includes moving the control wing to provide pitch, yaw and roll control.
According to yet another aspect of the present invention, a method of manufacturing a degradable missile control blade includes forming a frame from an organic resin reinforced with graphite and coating the frame with a thermal protection layer.

  According to one or more embodiments of the present invention, the step of forming the frame comprises high temperature resistance selected from the group comprising PMR-15 based polyimide, PT cyanate ester, bismarimide, phthalonitrile, abimide N. Using an organic resin, the step of coating the thermal protection layer includes laminating an ablation insulator to the frame, and the step of providing the thermal protection layer comprises chopped quartz, carbon, silica fiber Laminating one or more materials selected from the group comprising a phenolic resin reinforced with, wherein the step of providing a thermal protection layer includes a compression mold and the step of providing a thermal protection layer is continuous. Wrapping the fiber reinforced laminate to the frame and providing a thermal protection layer Forming a frame comprising mechanically interlocking the thermal protection layer to the frame by drilling a number of holes into the frame and molding into the hole to attach the thermal protection layer with additional adhesive. Includes attaching a metal shaft to the frame and / or forming a thermal protection layer includes forming a thermal protection layer over at least a portion of the shaft.

  One or more embodiments further include the step of attaching a corrosion protection material to the front edge of the frame, wherein the step of attaching the corrosion prevention material includes forming a carbon-carbon insert, the carbon-carbon insert being The forming step comprises a coating formed from a material selected from the group comprising a zirconium boride ceramic and a hafnium boride ceramic reinforced with a material selected from the group comprising silicon carbide and an ultra high temperature resistant ceramic. The step of providing a coating includes the use of a melted slurry process, and the step of providing a coating is typically less than about 1 / 5,000 inch (0.127 mm) thick. Forming a coating that is a step of providing a coating as a transitional tie layer. Including using an adhesive reactive sintered material, and the step of using a transitional bonding layer includes using a material selected from the group comprising hafnium carbide, hafnium boride, and corrosion protection The step of attaching the material further includes providing a coating selected from the group comprising an oxygen permeable barrier and a volatile barrier, wherein the step of providing a coating is selected from the group comprising a ceramic and oxide based coating. The step of forming the insert includes the use of a material selected from the group comprising a zirconium boride ceramic and a hafnium boride ceramic reinforced with silicon carbide fiber. And the step of forming the insert includes the use of ultra-high temperature resistant ceramic and forming the insert The step includes forming an insert having a substantially T-shaped cross section with a T tail connected to the frame, and / or forming the insert includes an open end of U attached to the frame. Forming an insert having a generally U-shaped cross section.

  According to yet another feature of the invention, the thrust vector control wing includes a frame for temporarily providing a structural support and a thermal protection laminate provided at least to the frame for temporarily thermally isolating the frame. And a combination of a corrosion prevention insert connected to the front edge of the frame to temporarily shield the frame from particle impact.

The foregoing and other features of the present invention will be described more fully hereinafter, particularly pointed out in the claims, the following description and the accompanying drawings, in which certain exemplary embodiments of the invention are described in detail. 1 illustrates one of the various ways in which the principle is used.
The present invention provides a thrust vector controlled jet wing, a jet wing control system, a missile with such a control wing (also referred to as a thrust wing or jet wing), and a method of manufacturing such a control wing. The control wing forms a control surface that generates the operating force of the vehicle by deflecting the rocket exhaust jet with missile launch. The method of manufacturing a composite control wing includes using both an ablation organic composite material and a refractory material. The control wing is able to obtain sufficient speed (called aerodynamic control speed), and when the aerodynamic control surface outside the fuselage controls the operation, it takes a certain amount of time to particles of particle size or dust size in the exhaust jet. Designed to be destroyed by. The jet wing control system provided by the present invention may be integrated into the original missile design or added to the missile as part of a kit or as a retrofit.

  An example missile 10 is shown in FIG. The missile has a fuselage or fuselage 12 containing a rocket motor and a missile payload (not shown). When the missile reaches the aerodynamic control speed, the missile also includes one or more aerodynamic control surfaces formed by wings or fins 14 projecting outward from the fuselage for controlled flight. The rocket motor is usually located in the direction of the rear or tail portion 16 of the missile and is thrust by an exhaust jet extending along the passage 18 to the rear end of the missile (right direction in FIG. 1).

  The jet vane control system is mounted in the rearward direction of the missile, usually (if any) outside the nozzle expansion cone so that the jet vane is located in the exhaust jet. As shown in FIGS. 2 and 3, the jet vane control system 20 includes a plurality of control vanes 22 that project inwardly in the direction of the exhaust jet passageway 18. The control system has, for example, three or four control blades that are equally spaced apart in a generally cylindrical direction around the passage of the exhaust jet. The jet vane control system further includes a control assembly 24 that typically includes one or more motors, bearings, and associated gears for controllably orienting the control vanes to a desired attitude with respect to the exhaust jet passage. Yes. A blast shield 26 is provided proximate the control wing to protect the components of the control assembly from the exhaust jet (may be formed integrally with the control wing as shown in FIG. 3).

  With reference to FIGS. 2-4, each control vane 22 includes an airfoil 30 that forms a control surface that interacts with the exhaust jet, and a control shaft 32 that connects the airfoil to the control assembly 24. For this control function, the shaft may be considered part of the control assembly. Both the control assembly (FIG. 2) and the control shaft control the direction of the airfoil of the exhaust jet. The control shaft is a synthetic resin shaft formed by a resin transfer mold manufacturing technique and a carbon fiber preform. Other materials with sufficient strength and temperature resistance properties such as corrosion resistant steel (CRES) or titanium-zirconium-molybdenum (TZM) or other refractory materials can be used in the shaft. Compared to synthetic resin shafts, refractory metal shafts provide greater heat resistance due to direct impingement of the exhaust jet if the shaft is not otherwise shielded. Blast shield 26 is required for plastic shafts and introduces unnecessary complexity if not otherwise required to protect internal control assembly components.

  Airfoil 30 may be formed around shaft 32, or the shaft may be attached after formation of the airfoil. For example, as shown in FIG. 4, the shaft includes a yoke 34 having dimensions for mounting the shaft to the airfoil base 36. The shaft can be glued to the airfoil or connected to the airfoil by fixing it to a mechanical attachment means by means of a mechanical fixture such as a nut and bolt combination, rivet or screw, for example. Good.

  An exemplary embodiment of the control vane 22 is shown in FIGS. 5A and 5C and 6A to 6C, and includes a control shaft 32 and an airfoil 30. FIG. The composite airfoil includes a frame 38, a leading edge insert 40 connected to the front edge 42 (rocket motor direction) of the frame, and a thermal protection system (TPS) 44 laminated to the frame.

  The composite frame of the airfoil 30 transmits drag, lift and hinge moment loads introduced by the exhaust jet to the control shaft 32, the control assembly 24 (FIG. 2), and the fuselage 12 (FIG. 1) of the missile 10. The frame is the primary load bearing structure of the airfoil and is formed of a high temperature composite such as a carbon reinforced high temperature organic resin, such as a PMR-15 based polyimide, PT cyanate ester, bismarimide, or Other advanced materials such as phthalonitrile, Avimid N, etc., but other similar high temperature materials are also suitable. In a few seconds, these materials can withstand an instantaneous thermal environment that exceeds about 1000 degrees Fahrenheit (about 538 ° C.), a typical temperature for ultrasonic missile applications. However, after being exposed to these high temperatures, these materials soften and oxidize, eventually disappearing as a structure and break down into granular particles.

  A refractory leading edge insert 40 is attached to the front edge 42 of the frame 38 to protect the composite frame and TPS 44 from direct impact of the rocket motor jet, especially from solid particles in the exhaust jet generated by the rocket motor propulsion. To do. The refractory leading edge is designed to withstand the initial thermal shock and load from the rocket motor ignition and propulsion flow until the aerodynamic control speed is reached. However, the leading edge is not designed to be durable so as to unnecessarily extend the life of the airfoil 30 beyond the time required to achieve this goal.

  7A and 7B show two possible structures for the leading edge insert 40, namely the plug structure of FIG. 7A and the shell structure of FIG. 7B. The shape and thickness of the leading edge is carefully selected to lose its effect in a predetermined manner and time period. These figures are merely illustrative and do not imply all possible structures.

  The leading edge is formed from a corrosion inhibiting material. If the leading edge 40 is made from a glass-, carbon-, or quartz-phenolic material, the leading edge is immediately catastrophic in the exhaust jet. Thus, the leading edge is formed of a carbon-carbon composite that is protected from high temperature oxidation by hafnium boride or zirconium boride ceramic, hafnium carbide or tantalum carbide or metal rhenium refractory coating. These coatings are performed, for example, by vapor deposition, plasma sprayed or melted slurry processes, and are typically less than about 1/5000 inch thick (less than about 0.127 mm), It is sufficient to temporarily protect the composite base layer.

  The carbon-carbon composite base layer of the leading edge 40 has hafnium carbide as a transition or bonding layer that allows other ceramic or oxide-based protective coatings to be provided on top of the transition layer as an oxygen permeable or volatile barrier. It may be coated with a sticky reactive sintered material such as hafnium boride. Reactive sintered coatings are formed by reaction between coatings, such as boride powder particles, base layer carbon, thereby providing a graded composition that adheres well to the carbon-carbon substrate with surfaces and penetrating layers having obscure interfaces. Forming and improving the bonding of the protective coating.

  Alternative leading edge ceramic materials include monolithic zirconium boride or hafnium boride ceramics reinforced with carbonized silicate fibers or particles. These ceramics are commonly known as ultra high temperature ceramics (UHTC) and have been developed for hypersonic aerospace vehicles such as sharp nose tips and leading edges. Carbon-carbon and UHTC composites can withstand high temperatures, including temperatures of about 4,000 to about 5,000 degrees Fahrenheit (about 2204 ° C. to about 2760 ° C.) without failure.

  With reference to FIGS. 5A-5C, as described above, a thermal protection system or TPS 44 may be laminated to the surface of the frame 38 and may also cover the leading edge 40. TPS provides ablation cooling and thermal isolation from propulsion aerial heat from the exhaust jet, thereby allowing the composite frame to temporarily perform its structural functions. TPS laminates are made from phenolic resin reinforced with cut quartz, carbon, and silica fibers, which are characterized as ablation insulators that serve for years as thermal shields for rocket motor nozzle components and re-entry vehicles Attached.

  TPS44 is formed by continuous fiber reinforced lamination of direct compression mold or tape wrapping to composite frame 38 to increase insulation uniformity, ablation performance, and ability of orthotropic materials . The TPS is mechanically interlocked to the composite frame by drilling a number of holes into the frame, filling the holes with TPS material, and curing the TPS material to solidify. After the holes are filled during the compression mold or filled with TPS material, the frame is wrapped with a TPS stack over the airfoil edge before the TPS is cured. These techniques allow the airfoil 30 to be accurately formed into any desired shape, including existing control wing shapes that are formed from non-degradable metallic materials.

  TPS materials usually do not have sufficient structural strength and robustness to function as the control wing 22 without the structural support provided by the frame 38 and usually corrode very quickly so that the control wing is fireproof. It does not make it possible to withstand the particle flow in the exhaust jet without the positive leading edge 40. However, the airfold 30 manufactured in accordance with the present invention uses a combination of these elements to form an excellent control wing due to the advantages described herein.

  The thrust control wing 22 formed in accordance with the present invention has a substantially repeatable and predictable failure mechanism designed to function within a predetermined time envelope of known stochastic tolerance. The control wing 22 and airfoil 30 formed in accordance with the present invention will not work by properly breaking down, i.e., breaking down into particles and / or dust sized particles within a predictable time period. Airfoil 30 typically does not include metal components because metal is generally more likely to be crushed and broken down into larger pieces rather than broken down into granular particles.

  The control wings 22 of the thrust wing control system 20 are designed to simultaneously disassemble to minimize the collapse of the missile flying along the intended trajectory. The failure of the composite airfoil 30 can be controlled by carefully selecting the composition of the airfoil subcomponents. The hottest organic composite has a glass transition temperature (Tg) of about 600 to about 800 degrees Fahrenheit (about 316 ° C. to about 427 ° C.), and continuous heat exposure does not cause material degradation. After the glass transition temperature is exceeded and the heat penetration time, all organic compounds release volatile materials and begin to oxidize. The stack of TPS 44 insulates the composite frame 38 to within the glass transition temperature during firing and pitch over, typically in less than a few seconds, after which large ablation degrades the TPS stack and its ability to cool the frame, The temperature begins to rise. Eventually, the TPS stack is consumed by ablation or delaminates from the release of the frame material when the frame material exceeds their glass transition temperature. In either method, once the composite frame is exposed to the exhaust jet, substantially instantaneous ignition and / or frame disassembly occurs. The refractory leading edge 40 similarly decomposes from the lack of structural support when the composite frame disassembles and from the inherent weakness of the refractory material used to form the leading edge.

  Depending on the specific manufacturer, the phenol TPS retraction rate typically exhibits temperatures in excess of about 4,000 degrees Fahrenheit (about 2204 ° C.), and in an exhaust jet environment containing a large amount of solid particles, about several thousand minutes per second. 1 inch (about 254 micrometers) or more. The larger the particle flow, the faster the corrosion of TPS44. In effect, rocket motors with a double grain propulsion design are specially configured to extrude two jets of different particle proportions. The first grain, for example, has a low particle flow to reduce the smoke on the surface ship and its sensor during missile launch and to ensure the residual force of the control wing 22 until aerodynamic control speed is reached. The second grain, starting immediately after the first grain is depleted, has a high particle content to increase the thrust for a special impulse and energized flight. The enhanced grain flow of the second grain and high stagnation temperature (delta T greater than about 5 hundred degrees Fahrenheit (about 260 ° C)) greatly increased the corrosion process of TPS44 and the decomposition process of the frame 38, and all control blades Increases the chance of destroying However, the initiated roll operation of the autopilot designed to overload all the thrust vector control composite frames, with a secured cruise flight trim, assuming that the TPS is approaching wear After the pitch over for control, simultaneous destruction of the jet wings can be facilitated.

  The secondary failure mechanism takes advantage of previously recognized shortcomings of using coated carbon-carbon composite components in rocket motor environments at stagnation temperatures in excess of about 6000 degrees Fahrenheit (about 3316 degrees Celsius). The control wing 22 can be designed as follows. Carbon-carbon composites are known to release volatile materials and break refractory coatings for about 10 to about 15 seconds after ignition of the rocket motor. Outgassing quickly causes catastrophic failure from rapidly propelled carbon-carbon oxidation and particle corrosion. Once the refractory leading edge 40 is crushed, the composite frame 38 and thermal protection system 44 are exposed to propellant impingement, resulting in rapid corrosion and crushing.

  Another secondary failure mechanism is triggered by the high thermal conductivity of the refractory material. Carbon-carbon and UHTC materials can withstand large amounts of thermal shock due to their low coefficient of thermal expansion and high thermal conductivity. Once the propellant impact is initiated, the refractory leading edge material “penetrates” thermal energy from the jet environment and conducts that heat to the composite frame 38. When the flame temperature exceeds the glass transition temperature, outgassing, oxidation, and loss of the structural support at the leading edge 40 cause rapid carbon-carbon or UHTC cracking and material failure. Once the leading edge breaks, the composite frame is exposed to direct propellant flow and the frame disintegrates quickly.

  When the timing is adjusted so that TPS 44 wears out just prior to crushing and disassembly of the secondary leading edge, the two mechanisms expose the composite frame 38 to the rocket jet environment and final evaporation. These fracture mechanisms are generated within the desired time envelope by new thermodynamic finite element modeling and retraction rate analysis and current material testing techniques when designing any thrust blade control system 20. Designed.

  The method of maneuvering the missile 10 (FIG. 1) during launch includes locating the thrust vector control wing 22 at the aft portion 16 of the missile so that the wing projects into the passage 18 of the rocket exhaust jet. The wing is connected to the steering control system 20 to allow controllable movement of the firing wing, thereby maneuvering the missile until aerodynamic control speed is reached. In operation, the missile is loaded onto a substantially vertical canister for launch. When a threat is recognized and the rocket engine is ignited for missile launch, the missile will controllably change the direction of the control wing 22 in the rocket exhaust jet for full roll, pitch, and yaw operation. It can be operated immediately upon launch. After a few seconds, the missile 10 pitches over and achieves aerodynamic control speed, and the thrust vector control wing breaks down with the exhaust jet, thereby reducing missile weight and improving thrust control and trim. The missile then searches for and destroys the target in the usual way.

  In summary, the decomposable thrust vector control jet wing 22 uses a high temperature composite frame 38 that forms a primary load carrying structure. An ablation organic synthesis thermal protection system (TPS) 44 is laminated to the composite frame for temporary thermal isolation from propulsion air heating. A leading edge insert 40 that joins the front edge 42 of the composite frame provides temporary protection from direct particle impact in the TPS and the corroding rocket motor nozzle environment on the frame. After erosion of the leading edge, the composite control wing is typically placed in the granular particles in the exhaust jet after the TPS has been consumed in a predetermined manner in time based in part on the thickness of the provided TPS stack. Disassembled into

  Although the present invention has been shown and described with respect to the illustrated embodiments, equivalent changes and modifications can be made by one skilled in the art upon reading and understanding the specification and the accompanying drawings. In particular, with respect to the various functions performed by the aforementioned complete bodies (components, assemblies, devices, composites, etc.), they are not structurally equivalent to the disclosed structures performing functions in the presently illustrated embodiments of the present invention. However, the terms used to describe such completes (including references to “means”) correspond to any completes that perform a specific function (ie functionally equivalent), otherwise It is shown in the method. Further, while the special features of the present invention have been described above with respect to only one of several illustrated embodiments, such features may be useful when desired and effective in any given or special application. It may be combined with one or more other features of the embodiments.

1 is a front view of a missile having a jet wing control system attached to the aft section. FIG. FIG. 3 is a perspective view schematically showing a part of a jet blade control system partially cut away. FIG. 3 is an enlarged perspective view of the control wing and blast shield from the control system shown in FIG. 2. The schematic exploded view of the control wing which shows the method of attaching a control shaft to a wing (airfoil). The side view of the control wing | blade according to this invention, the top view of the control wing | blade, and sectional drawing when it sees along the line 5C-5C of the top view. Side view of the airfoil portion of the control wing in accordance with the present invention without the wing (airfoil) thermal protection system portion, cross-sectional view when viewed along line 6B-6B of the side view, and side view thereof Sectional drawing when it sees along line 6C-6C. An enlarged partial cross-sectional view of a control wing according to the present invention showing one embodiment of the leading edge portion of the wing (airfoil) and an enlarged view of the control wing showing an alternative embodiment of the leading edge portion FIG.

Claims (9)

In the decomposable thrust vector control wing ( 22/30 ) mounted in the rocket nozzle to be located in the exhaust jet passage from the rocket motor ,
A movably configured frame (38) made of composite material and protruding into the nozzle ;
Thermally protective layer provided on at least on part of said frame (38) and (44),
Provided on the front edge (42) of said frame (38), said frame (38) before being configured with a corrosion inhibiting material to protect the edge insertion portion with respect to the solid particles contained in the exhaust gas jet (40) equipped with,
The corrosion prevention material of the leading edge insert ( 40 ) and the material of the thermal protection layer ( 44 ) are selected to withstand the conditions in the exhaust jet for a predetermined length of time,
The frame ( 38 ) is a decomposable thrust vector control blade (22/30) configured to be decomposed into granular particles in the exhaust jet after the elapse of the predetermined length of time . The corrosion preventing material carbon - carbon structure and ceramic, carbide, and degradable control vane according to claim 1, wherein and a surface coating on the selected structure from the group comprising a metallic material. 3. The decomposable control blade according to claim 1, wherein the frame (38) contains an organic resin reinforced with graphite. The organic resin reinforced with graphite was reinforced with carbon containing one or more resins selected from the group comprising PMR-15 based polyimide, PT cyanate ester, bismarimide, phthalonitrile, abimide N. The decomposable control blade according to claim 3, which is a high-temperature organic resin. Thermally protective layer (44) is degradable control vane according to any one of claims 1 to 4 includes an ablation insulator. 6. The degradable control wing of claim 5 , wherein the thermal protection layer (44) comprises one or more materials selected from the group comprising phenolic resin reinforced with chopped quartz, carbon, and silica fibers. In a thrust vector control system (20) for a missile (10) having a motor for propelling the missile ( 10 ) that generates an exhaust jet,
And control vanes of any one of claims 1 to 6 (22/30), a control assembly for controlling the direction of the control vanes includes a (2 4) and the control system (20) is controlled A thrust vector control system (20) configured to be attached to the missile such that the wings ( 22/30) are in the passage (18) of the exhaust jet. A motor for propelling a missile that generates an exhaust jet, and a control blade (22/30) according to any one of claims 1 to 6 mounted in an exhaust jet passage (18). Missile (10). In the method of maneuvering the missile (10) during the launch period ,
Degradable movable control vane of any one of claims 1 to 6 (22/30) disposed in the path of the exhaust jet of the missile (18), and launch missiles, said to steer the missile moving by controlling the control vanes (22/30), out injection exhaust jet by the firing of the missile, Ru to decompose the said control vanes (22/30) in the exhaust jet within a predetermined length of time steps Including methods.

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